63results about How to "Ensure flight stability" patented technology

A dual, coaxial rotor helicopter is provided that is relatively easy to fly. Thrust is provided by two ducted fans that are mounted at the rear of the aircraft and spaced apart laterally. Differential thrust generated by the fans provides yaw control for the aircraft, and forward thrust is provided by the fans working in combination. The coaxial rotors are preferably utilized primarily for lift, and not for forward thrust, which simplifies the control requirements. The coaxial rotor with ducted fan configuration also results in lower vibratory loads being imposed on the helicopter, thereby increasing its speed capability. The fan ducts serve to protect the fans, augment the fan thrust at low airspeeds, increase the efficiency of the fans at cruise speeds, and provide horizontal and vertical stabilizing surfaces to ensure aircraft flight stability.

The invention discloses a fault reconstruction method for a simple redundancy flight control system of a long-endurance unmanned aerial vehicle. The redundancy is formed for the measurement of key navigation information by using two devices of main and backup sensors on the basis of the non-similarity principle, and two redundancies are guaranteed on an execution mechanism, so that the simple redundancy flight control system is formed. A comparison is made between information of the main sensor and the information of the backup sensor; when the times of the deviation of the information of the main and backup sensors greater than the threshold in unit time are more frequency, one sensor is broken down; the fusion is implemented through other normal navigation information for locating the broken device; and when one execution mechanism is broken down, the influence of the broken mechanism is complemented through the rest normal execution mechanisms to guarantee the flight safety.

The invention discloses a single-drive synchronous unfolding mechanism provided with multiple folding control planes. Each control plane of a traditional small guided missile is independently driven by a torsion spring to be unfolded, synchronism is difficultly ensured, and the adopted unfolding mechanism enable the radial size of the guided missile to be large during folding. The single-drive synchronous unfolding mechanism comprises a cabin body, an actuator cylinder, a connecting rod and four unfolding mechanisms, wherein an actuator cylinder installing seat is fixed to the cabin, the actuator cylinder is connected with an actuator cylinder installing seat sliding pair, four actuator cylinder convex lugs are evenly distributed on the actuator cylinder, each actuator cylinder convex lug is hinged to the inner end of one connecting rod, and the outer ends of the four connecting rods extend out of installing holes formed in the unfolding mechanism respectively and are hinged to slide blocks of the corresponding unfolding mechanism. The unfolding mechanism comprises a rudder stick, the slide blocks, lug pieces, a middle rudder, a rudder tip, connecting rods and a spring locking mechanism, wherein the spring locking mechanism is used for locking the slide blocks and a rudder shaft. The transverse size of a missile body is effectively decreased during control plane folding. The controllability and synchronism in the unfolding process of the four control planes are ensured through driving of the single actuator cylinder.

The invention provides an Opencv-image-processing-based quadrotor aircraft midair autonomous grabbing operation control method. A quadrotor aircraft operation system is composed of a quadrotor aircraft and an operation device, and the operation device is composed of a mechanical arm and a gravity center balance mechanism, wherein the mechanical arm and the gravity center balance mechanism are symmetrically installed below the quadrotor aircraft. Images of a grabbed target and an artificial mark are obtained through a camera on the quadrotor aircraft, the Opencv is utilized for processing the images, the camera posture information is solved, and thus the posture information of the quadrotor aircraft is obtained; and the quadrotor aircraft, the mechanical arm and the gravity center balance mechanism are controlled in a graded manner according to the posture information of the quadrotor aircraft. The application filed of the quadrotor aircraft is broadened, midair operation can be performed either in a room or in a surrounding outdoor environment with dense buildings, and midair operation precision is improved.

The invention provides an unknown input observer based realization method of fault diagnosis for a minitype unmanned aerial vehicle distributed formation. According to a hierarchical concept of a distributed control system, controller designs of the single unmanned aerial vehicle and multi unmanned aerial vehicle formation are studied to ensure the flight stability of the minitype unmanned aerialvehicle distributed formation. When a fault of a single unmanned aerial vehicle final controlling element occurs, an effective unknown input observer distributed system fault detection method is provided, which completely decouples interference and completely eliminates adverse effects of the external interference on a fault diagnosis process. Furthermore, a method to estimate the fault by extending a state vector and a fault vector into an augmented vector is provided. According to the invention, the conventional unknown input observer is combined with an augmented system to realize fault detection, separation and estimation of the minitype unmanned aerial vehicle distributed formation. According to known minitype unmanned aerial vehicle model parameters, a numerical simulation can be established to carry out fault diagnosis on the final controlling element. The fault diagnosis method is used for fault diagnosis of a minitype unmanned aerial vehicle distributed formation.

The invention discloses a composite aircraft composed of fixed wings and deformable electric multi-rotary wings. The composite aircraft comprises a machine head, a machine body, a tail wing and wings. Pushing propellers are arranged in the front of the machine head. Rotary wing rods are connected with the wings through rotation devices with drive devices. The upper end faces of the rotary wing rods are ridge-shaped streamline hook surfaces with the two sides arching towards the middle, and the rotary wing rods are of a telescopic rod structure. The composite aircraft has the beneficial effect that vertical taking-off and landing of a rotary-wing helicopter are not limited by a taking-off and landing field and also has the flying speed and durable cruising capability of a fixed-wing airplane. The rotary wing rods can be bent, overturned and rotated through cooperation of bending devices, rotating devices and roller devices, and therefore the propellers can be controlled by a master controller to change the direction and angle at will, and flying is more flexible and convenient. The invention further discloses a control method of the composite aircraft composed of the fixed wings and the deformable electric multi-rotary wings

The invention discloses a submerged folding wing synchronous lateral unfolding locking mechanism and an unfolding locking method thereof. Existing airfoil folding and unfolding locking mechanisms areexposed outside projectiles in the folded state, the overall aerodynamic performance of missiles is affected in the folded state, and the folding and unfolding process is greatly affected by fluid. The mechanism of the invention comprises a machine frame, folding wings, a spring locking mechanism, a positioning block, a motor, a locking sliding block, a pin shaft, a screw rod, a torsional spring fixing member, a torsional spring, a nut block, a coupling, a spring and a spring pin. The N folding wings are synchronously unlocked by moving the locking sliding block in the axial direction of the screw rod; and the unlocked folding wings are unfolded under the action of the torsional spring, and are locked by the spring locking mechanism after being unfolded in position. The mechanism is storedin a projectile in the folded state, submerged folding of the missile wings is achieved, the aerodynamic performance of the missile is effectively ensured, and the lateral dimension of the missile isreduced.

The invention discloses a flight control method and system for an unmanned aerial vehicle, and solves the problem that the flight stability and the security of the unmanned aerial vehicle cannot be ensured as only one parameter of air speed or ground speed is used for flight control of the unmanned aerial vehicle. The flight control method and system for the unmanned aerial vehicle timely detects the air speed and the ground speed of the unmanned aerial vehicle and timely corrects to obtain the real time true air speed and ground speed of the unmanned aerial vehicle, which reflect the speed of the unmanned aerial vehicle relative to the air in a more real and objective manner. Based on this, the flight control of the unmanned aerial vehicle is carried out with the combination of the real-time real air speed and the real-time real ground speed, so that the flight stability and security can be effectively ensured when the unmanned aerial vehicle is influenced by factors such as turbulence flow, wind shear, crosswind, low level ground effect and the like.

The invention discloses an amphibious aircraft with an automatic balancing empennage. The amphibious aircraft comprises an aircraft body and an engine installed on the aircraft body. An air propeller is connected to an output shaft of the engine and located on the upper side of the back portion of the aircraft body, and the horizontal empennage is arranged at the portion, in slip flow generated by the air propeller, of the upper side of the back portion of the aircraft body; on one hand, the attack angle of the horizontal empennage is adjusted to offset the angle of the negative pitching moment of the engine in the take-off stage; on the other hand, the horizontal empennage serves as an adjusting sheet for the pitching balance of the aircraft in the flight process. In the take-off stage, the attack angle of the horizontal empennage is adjusted to offset the angle of the negative pitching moment of the engine; in this way, even when the aircraft is at the critical take-off speed, the engine is suddenly shut down, the pitch attitude of the aircraft cannot be changed, and the flight safety of a pilot is guaranteed accordingly. When the aircraft normally flies in the air, the horizontal empennage is adjusted to be at the proper angle and serves as the adjusting sheet for the pitching balance of the aircraft to achieve the attitude trimming function, and the flight stability of the aircraft is guaranteed.

The invention relates to a parachute system for aircraft recovery and a parachute recovery method, and belongs to the technical field of rocket recovery. When a drag parachute is opened, a hanging point is positioned behind the mass center of an aircraft, so that the stability of the attitude of the aircraft is favorably kept, and the drag parachute is in close contact with the wall surface of a main parachute bay due to the large size of the main parachute; if the exit angle of a main parachute bag is smaller, the main parachute bag is easily clamped in a main parachute cabin to cause failureof opening the main parachute, so that the hanging point of the drag parachute is instantly switched to the front of the aircraft before the main parachute bag exits the cabin, the aircraft is liftedup, a certain angle is created for the exit of the main parachute bag, and smooth exit of the main parachute bag is realized; when the main parachute is opened, the opening force is larger, and if the aircraft with a single hanging point structure is adopted, the aircraft is easy to roll and is wound with a hanging belt of the main parachute, the parachute opening failure is caused, and the horizontal recovery is difficult to realize; and a front hanging point support of the main parachute is fixed in front of the center of mass, and a two rear hanging point supports are fixed behind the center of mass, so that the aircraft can be prevented from being overturned and wound when the main parachute is opened, and the horizontal recovery of the aircraft can be realized.

The invention belongs to the technical field of fire extinguishing equipment, in particular to a fire-extinguishing grenade 120 and a launcher used with the same, which are used for solving the problems that existing forest fire-fighting grenades are single in application range and not powerful enough and launchers for the existing forest fire-fighting grenades are unable to meet the modern fire extinguishing requirements. The caliber of a bullet of the fire-extinguishing grenade 120 is 120mm, and a ventilating bourrelet and a bourrelet of the fire-extinguishing grenade 120 are 120mm in largest outer diameter. A plurality of vent grooves are evenly and axially distributed on the outer circular surface of the ventilating bourrelet. A plurality of belts are symmetrically disposed on the outer circular surface of a belt block, each belt has a twining angle of 6 DEG and 24''. A powder chamber is in a concave chamber structure. The launcher comprises a body pipe connected with the tail of a mortar shell 60 through a connector, the body pipe is 120mm in inner diameter, and rifling grooves corresponding to the belt block are symmetrically reserved in the body pipe. The fire-extinguishing grenade 120 and the launcher used with the same have the advantages that the method of launching the grenade as launching the mortar shell is realized, fire extinguishing methods are expanded, and the fire-extinguishing effect is improved.

The invention discloses a cable erecting system based on an unmanned aerial vehicle and a working method thereof and belongs to the technical field of cable erecting. The system comprises the unmannedaerial vehicle, a tension machine and a motorized winch, wherein a traction rope winding and unwinding mechanism is arranged at a lower end of the unmanned aerial vehicle, the traction rope winding and unwinding mechanism comprises a winding and unwinding assembly, a tension detection assembly, an adjusting assembly and a first traction rope, the tension machine is provided with an installation cable, one end of the installation cable is provided with a second traction rope, one end of the second traction rope is fixedly connected with the installation cable, the other end of the second traction rope is matched with a free end of the first traction rope, and the motorized winch is matched with the second traction rope and the installation cable. The system is advantaged in that by arranging the tension detection assembly, the tension of the first traction rope can be detected, flight stability of the unmanned aerial vehicle is guaranteed, and by arranging the first traction rope and the second traction rope, the situation that the traction ropes are broken in the working process can be avoided.

The invention relates to a spin-stabilized falling-area precise control system for rocket separator remains. The spin-stabilized falling-area precise control system for rocket separator remains comprises a spinning system, a grid fin system and a separating body, wherein the separating body comprises a separating body head and a separating body; one end, close to the separating body head, of the separating body is equipped with a spinning engine system, and the tail part of the other end is equipped with the grid fin system; the spinning engine system comprises two spinning engines; and spray tubes of the two spinning engines are distributed in an axisymmetric mode along the axis of the separating body. Two symmetrical engines are adopted for spinning a rocket, so that higher spinning rotation speed can be obtained, influences of interference factors such as pneumatic asymmetry and process errors can be effectively reduced, and flying stability of an object is kept; and a combined control way of spin-stabilizing and the grid fin, so that the falling-area range of the rocket remains can be effectively reduced to 100 square kilometers (10 km*10km), and adverse effect on an emitting task of the falling-area problem can be reduced and even avoided.

The invention discloses a solar aircraft and a control method of a controller of the solar aircraft.The solar aircraft comprises a body, thin-film solar cells and the controller, wherein the thin-film solar cells are arranged on the surface of the body, and the controller is arranged in the body.The body comprises two symmetrical flying wings, each flying wing at least comprises a first end wing and a second end wing, and each first end wing and the corresponding second eng wing are rotatably connected through a first rotating shaft.The controller is used for controlling rotation of each first rotating shaft according to sunlight irradiation angles so as to increase radiation areas of the thin-film solar cells arranged on the surface of the body and solar radiation energy received in unit area.Therefore, endurance time of the solar aircraft is prolonged, and flying reliability of the solar aircraft is improved.

A sewn ball includes an inner tube and a housing. The inner tube includes a tube body, a yarn layer and a rubber layer, wherein the yarn layer is attached to an external surface of the tube body, the rubber layer is attached to an external surface of the yarn layer and is then vulcanized, in such a manner that a rubber raw material of the rubber layer permeates through yarn gaps of the yarn layer, so as to integrate the tube body, the yarn layer with the rubber layer. The housing wraps around and is attached to the inner tube. The sewn ball provided by the present invention has excellent jumping ability, high strength, good durability, small circumferential error, and excellent balance and flight stability.

The invention discloses an unmanned aerial vehicle for eliminating power current overhead lines and an eliminating method thereof. The unmanned aerial vehicle for eliminating the power current overhead lines comprises a bottom plate, a stand column and a frame, wherein the frame comprises a middle flat plate part and four wing parts arranged around the middle flat plate part, and motor supports are arranged at the upper parts of the ends of the four wing parts; a motor is arranged on each motor support, and a rotor wing is connected onto an output shaft of the motor; two mechanical arms are arranged on the two sides of the bottom of the bottom plate, a mechanical claw is connected onto each mechanical arm; the bottom of the bottom plate is connected with a camera cradle head, and a camerais connected onto the camera cradle head; two landing gears are arranged on the two sides of the bottom of the bottom plate, and a control console is arranged at the top of the bottom plate; and a control circuit is integrated in the control console and comprises a microcontroller module, a power module, a wireless communication module, an infrared receiving module, an attitude sensor and four motor driving modules. The unmanned aerial vehicle has the advantages of novel and reasonable design, convenience in realization, high safety, high eliminating efficiency, low cost of human and materialresources, high practicality and convenience in promotion and application.

The invention discloses a rotor-free unmanned aerial vehicle in the technical field of unmanned aerial vehicles, the rotor-free unmanned aerial vehicle comprises a vehicle body and four sets of vehicle arms annularly arranged on the periphery of the vehicle body, an air inlet bin is arranged in the vehicle body, the air inlet bin is provided with air inlets communicating to the two sides of the vehicle body, and the vehicle arms are each of a hollow pipe body structure; the inner side ends of the machine arms are communicated with the air inlet bin through soft air pipes, the outer side ends of the machine arms are communicated with airflow jet rings, turbofans driven by servo motors are arranged in the four sets of machine arms respectively, and a battery mainboard bin is arranged at the position, located below the air inlet bin, in the machine body. According to the rotor-free unmanned aerial vehicle, most topographic conditions can be ignored, the unmanned aerial vehicle is different from an unmanned aerial vehicle with external rotor wings, blades are prone to being damaged during flight, the unmanned aerial vehicle is damaged, and the unmanned aerial vehicle with the built-in rotor wings can be ignored from the dangerous factors, so that the unmanned aerial vehicle with the built-in rotor wings is safe and reliable. And the unmanned aerial vehicle has good flight stability, controllability and safety.

The invention discloses an all-terrain photovoltaic cleaning system based on a large-load unmanned aerial vehicle and a working method of the all-terrain photovoltaic cleaning system, and belongs to the technical field of photovoltaic power generation. The system comprises an unmanned aerial vehicle, a water tank, a water supply system, a nozzle assembly, a power supply system, a visible camera, an infrared camera and a control system. The water tank is detachably connected to the bottom of the unmanned aerial vehicle; one end of the water supply system is connected with the water tank; the power supply system, the visible light camera and the infrared camera are arranged below the unmanned aerial vehicle; the power supply system is connected with the unmanned aerial vehicle, the water supply system, the visible light camera, the infrared camera and the control system. The control system is connected with the unmanned aerial vehicle, the water supply system, the visible light camera and the infrared camera. The method is good in terrain compatibility and high in efficiency, saves the cost, and saves manpower and material resources.

The invention discloses a rotation assembly. The rotation assembly comprises a rotation angle piece, an angle positioning piece rotating relative to the rotation angle piece, and a hinge locking mechanism acting between a locking position and an unlocking position. At least one clamping part is arranged on the rotation angle piece in the rotating circumferential direction at intervals, a positioning part matched with the clamping part is arranged on the angle positioning piece, and when the hinge locking mechanism is located at the locking position, the positioning part and the clamping part are kept in a matched state. The rotation assembly is simple and compact in structure, good in rotating stability, high in precision and capable of reliably locking the rotating state, that is, the aircraft is kept in the unfolded state and the folded state stably without displacement, so that the flight stability and safety are guaranteed, the stability of the folded state is guaranteed, and the safety and reliability of the rotation assembly in the placing and transporting process are guaranteed.

The invention discloses an aircraft for geographic surveying and mapping in a plateau area. The aircraft comprises an unmanned aerial vehicle, an external carrying box arranged at a bottom of the unmanned aerial vehicle and a ground terminal. The unmanned aerial vehicle is connected with the external carrying box through a communication cable, and the unmanned aerial vehicle and the external carrying box are in wireless connection with the ground terminal. In the present invention, the aircraft for geographic surveying and mapping in the plateau area can acquire an image of the plateau regionand detect water pollution, plant production change, mineral products, lands and other resources through a camera and a remote sensing surveying and mapping module and detect a meteorological cloud layer, functions are complete, an image processing module is arranged in a matched mode, data compression, image recovery, information extraction and image enhancement can be conducted on the image, better storage and transmission are facilitated, the improvement of the data processing efficiency is facilitated, the statistical analysis, automatic recognition, classification and integration of information are achieved and the visual quality is improved, and a user can obtain targeted information.

The invention relates to a parachute system for aircraft recovery and a parachute recovery method, and belongs to the technical field of rocket recovery. When a drag parachute is opened, a hanging point is positioned behind the mass center of an aircraft, so that the stability of the attitude of the aircraft is favorably kept, and the drag parachute is in close contact with the wall surface of a main parachute bay due to the large size of the main parachute; if the exit angle of a main parachute bag is smaller, the main parachute bag is easily clamped in a main parachute cabin to cause failureof opening the main parachute, so that the hanging point of the drag parachute is instantly switched to the front of the aircraft before the main parachute bag exits the cabin, the aircraft is liftedup, a certain angle is created for the exit of the main parachute bag, and smooth exit of the main parachute bag is realized; when the main parachute is opened, the opening force is larger, and if the aircraft with a single hanging point structure is adopted, the aircraft is easy to roll and is wound with a hanging belt of the main parachute, the parachute opening failure is caused, and the horizontal recovery is difficult to realize; and a front hanging point support of the main parachute is fixed in front of the center of mass, and a two rear hanging point supports are fixed behind the center of mass, so that the aircraft can be prevented from being overturned and wound when the main parachute is opened, and the horizontal recovery of the aircraft can be realized.

The invention relates to an aircraft airfoil surface structure which is composed of a tree-shaped metal inner core and a composite material skin, in particular to a structure of a rudder wing surfaceat the tail part of a hypersonic aircraft used under a high temperature service environment. The airfoil surface structure comprises the inner core, the skin, a rudder shaft and connection bolts; theinner core is designed to be of a tree-like I-shaped beam structure and prepared from a GH4099 high temperature alloy, and the inner core has the excellent mechanical property under the high temperature environment; the skin is prepared from a C / SiC composite material, and is good in heat resistance and density; the inner core and the skin are connected with the aircraft through the rudder shaft,and the rudder shaft is prepared from the GH4099 high temperature alloy; and the inner core and the skin are connected through ceramic bolts. The airfoil structure can guarantee the mechanical property and the heat resistant property of the aircraft under the high temperature and hypersonic flying environment, and improves the flight stability.

The invention aims to provide a wing interior supporting structure of an unmanned aerial vehicle. The wing interior supporting structure of the unmanned aerial vehicle comprises a front transverse beam, a rear transverse beam, a connecting baffle and a plurality of supporting plates. The front transverse beam comprises a front transverse beam outer rod and a front transverse beam inner rod, and the rear transverse beam comprises a rear transverse beam outer rod and a rear transverse beam inner rod. The front transverse beam inner rod comprises a front transverse beam inner rod metal rod and a front transverse beam inner rod composite rod. The rear transverse beam inner rod comprises a rear transverse beam inner rod metal rod and a rear transverse beam inner rod composite rod. The front transverse beam and the rear transverse beam are arranged in parallel. Any one supporting plate is perpendicular to the front transverse beam and is movably connected with the front transverse beam. Any one supporting plate is movably connected with the rear transverse beam. The connecting baffle is provided with a front connecting through hole and a rear connecting through hole. The wing interior supporting structure of the unmanned aerial vehicle has the beneficial effects that by means of the two transverse beam structures each of which is made of metal and a composite material in combination, vibration of a wing interior structure in the flying process of the unmanned aerial vehicle is reduced and the flying stability and flying speed are guaranteed on the premise that the strength is guaranteed and the whole wing weight is not markedly increased.

The invention discloses a dynamic force bearing center position detection method for a symmetrical-layout multi-rotor unmanned aerial vehicle. The method comprises steps: (1) when the unmanned aerial vehicle is motionless, the geometric central position information of a rotor surface is acquired, and a geometric central position graph equation is obtained; (2) in the case of flying when the unmanned aerial vehicle takes off, the graph equation is used to obtain gravity center position information of the unmanned aerial vehicle, and the information is recorded in a dynamic vector space coordinate point form; (3) the rotation speed and the spatial attitude position information of each rotor of the unmanned aerial vehicle are acquired, data features of the two are extracted, and a flying path equation and a spatial attitude equation for the unmanned aerial vehicle are obtained; (4) the spatial attitude equation is used to be combined with the geometric central position graph equation, a dynamic force bearing center position equation for the unmanned aerial vehicle is obtained; (5) the dynamic spatial vector coordinates of the gravity center and the dynamic spatial vector coordinates of the dynamic force bearing center position are compared in real time, and the dynamic change relationship between the two is obtained; and (6) a flying control system processes and compares the dynamic relationship between the two in real time, and the unmanned aerial vehicle is assisted for attitude adjustment.

The invention provides a composite rotor aircraft and relates to the aircraft technology. According to the composite rotor aircraft, by arranging an aircraft body and wings which are fused together, the exposed adverse soaking area of the aircraft body can be reduced and the aerodynamic resistance of the composite rotor aircraft can be reduced; and by mounting an integrated flaperon, the flying time of the composite rotor aircraft can be prolonged, the flying distance of the composite rotor aircraft can be extended and the loading capacity of the composite rotor aircraft can also be enhanced;and moreover, after a takeoff and landing motor power system is detached, the composite rotor aircraft is suitable for being remodeled into a fixed aircraft fused together with the wing bodies. Meanwhile, by detachably connecting a tail wing assembly with a supporting rod, the tail wing assembly can also be detached in the process of transporting, grounding or so on so that the tail wing assemblyis prevented from being knocked and the working reliability of the tail wing assembly is guaranteed.

The invention discloses an unmanned aerial vehicle bridge floor coverage motion planning method, belongs to the technical field of unmanned aerial vehicle control, and aims to solve the motion planning problem that an unmanned aerial vehicle stably completes the full coverage of the bridge floor and meets the coverage resolution requirement in an automatic detection task of a bridge surface. The method comprises the following steps of firstly, establishing a geographic coordinate system, an airframe coordinate system and a bridge floor coordinate system, giving a measurement model of an airborne camera, and giving the definitions of a coverage space, a coverage surface and the coverage resolution; secondly, decomposing a topological structure of the bridge floor by adopting a linear scanning form, and determining a bow-shaped coverage path for each decomposition unit; then determining the coverage sequence of the units by adopting an integer optimization method; then giving two mathematical descriptions of the flight space of the unmanned aerial vehicle under a coverage task by combining the distance range from the unmanned aerial vehicle to the bridge floor; and finally, taking the flight stability of the unmanned aerial vehicle as an index, and combining with limitations, such as position, speed, acceleration, flight space, full coverage requirements, etc., providing a method for obtaining a motion trail function by adopting curve optimization. The method is suitable for the unmanned aerial vehicle to quickly and stably complete the movement track planning of full coverage of the bridge floor under a bridge detection task.

A disclosed water quality monitoring unmanned aerial vehicle capable of maintaining dynamic balance comprises a vehicle body, a working cavity is formed in the vehicle body, six first connecting rods are evenly and fixedly connected to the circumferential side face of the vehicle body, and a rotor wing assembly is fixedly connected to the end, away from the vertical center line of the vehicle body, of each first connecting rod. A sampling device for sampling and memorizing a water body to keep the stability of the unmanned aerial vehicle is arranged on the lower side of the vehicle body. A supporting rod and a third connecting rod are arranged, so that when the sampling device moves downwards, the third connecting rod swings left and right, and the third connecting rod drives the supporting rod to swing left and right through a second connecting rod; and floating objects on the water surface are pulled away, so that the sampling is prevented from being hindered.

A double-cylinder water cooling power device for an unmanned aerial vehicle comprises an engine component and a power generator component; the engine component comprises a first cylinder and a secondcylinder which are arranged at the two ends; cylinder body cooling boxes are correspondingly arranged on the external side of the first cylinder and the external side of the second cylinder; a circulation cooling component is connected with the cylinder body cooling boxes through cooling pipelines to form a cooling loop; and a cooling liquid flows in the cooling loop in circulation to cool the engine component. By using a water cooling circulation type cooling mode, a cooling speed of an engine can be increased; in particular for a plant protection unmanned aerial vehicle, a double-cylinder engine can be sufficiently cooled by using the circulation cooling component on the premise of not increasing the mounting space; and by combining natural wind power in the unmanned aerial vehicle flying process, a power generator is cooled so that thermal accumulation can be effectively avoided.

The invention provides two manned vertical take-off and landing and fixed-wing hovercars and two vertical take-off and landing and fixed-wing hovercars with manned and wireless remote control functions. A multi-layer continuous airfoil propeller is adopted to provide lifting force in the vertical direction, and the spiral airfoil structure of the multi-layer continuous airfoil propeller is utilized to cut and adapt to windward rapid laminar flow or crosswind laminar flow, so that the phenomena that the windward laminar flow or the crosswind laminar flow directly generates airflow impact on the blades and the wingtips of the blades generate shock wave vibration when a common blade type propeller is adopted, resulting in large wind resistance, unstable flight or hovering and yawing caused by wind blowing are avoided; a flight driving device is adopted to provide level flight power after lift-off, and during rapid level flight, after operation of a multi-layer continuous airfoil propeller is stopped, lift force is generated by a main wing to maintain level flight, so that fuel or electric power is saved, and the endurance time is prolonged; the attitude of each hovercar is adjusted by wing control surface actions, so that the flight control difficulty is reduced; the combined take-off and landing modes such as vertical take-off and landing and gliding take-off and landing can be achieved, and the adaptability to take-off and landing sites is improved.